1. Field of the Invention
The present invention generally to improvements in space vehicles and more particularly, but not by way of limitation, to a method and arrangement for providing a strain compatible attachment for a metal thermal protection system for a space vehicle and similar transatmospheric and high speed aerospace vehicles.
2. Description of the Prior Art
Future space vehicles in the heavy launch arena, such as reusable launch vehicles (RLV's), will be required to realize increased performance through the incorporation of durable, robust, low density, high stiffness thermal protection systems (TPS). Such TPS for an aerospace vehicle will require aircraft like operations and maintenance support to allow the next generation of RLV's to be economically feasible. In addition, such TPS must provide an effective means of transmitting aerodynamic loads (shear and pressure differences). Thus, the load transfer system must have low weight, result in low thermal stresses in the underlying aerostructure, and not transfer excessive heat to such underlying aerostructure.
Insulation structures, which consist of a plurality of highly heat-resistant panels to be connected to the airframe of a space vehicle have been known from the U. S. Space Shuttle Program, for example. During reentry into the atmosphere, individual panels or even panel fields often are damaged by thermal and flow mechanical effects so significantly that they must be replaced before the next launch. This replacement process is very time consuming and expensive due to the types of fastening for the TPS panels that have been used to date, e.g., adhesive bonding. A side from such criteria of replaceablility and repairability, it should be borne in mind that the fastening of the TPS panels to an airframe must be extremely robust and reliable in order to withstand the high flow mechanical loads occurring during certain phases of flight. On the other hand, however, the arrangement for attachment of the TPS panels should allow for thermal expansion and contraction of the panels, without generating undesired forces and stresses in the structure. In addition, fuel tank pressurization and flight induced mechanical deflections can not be allowed to interfere with TPS integrity and sealing capabilities. Also, the best possible protection of the fastening elements against high thermal loads is desirable.
One approach to satisfy these load bearing and thermal isolation functions at low thermal stresses employs a load bearing insulation attached to the primary aerostructure. Since most load bearing insulations have low strength and no ductility, they are usually segmented with small gaps for low thermal stress. One load bearing insulation previously developed includes the use of sintered quartz-fiber tiles to provide a reusable surface insulation for a RLV. While nonmetallic insulations are simple to attach to the aerostructure, they have a common potential disadvantage. That is they are weak, brittle materials, consequently, surface frayings, erosion, cracking or breakage rates may be high and may increase refurbishment requirements.
One fastening scheme for connecting each TPS panel to the next panel and to the airframe of a RLV is seen in U.S. Pat. No. 4,344,591. In this arrangement, each panel has, on its front side, two projecting tongues, which engage straps on the airframe as well as in recesses of the next panel. Thus, it is possible to install rows of panels in the circumferential direction of the fuselage of the space plane, wherein the first and last panels require a separate fastening on one side. However, it is obvious that no individual panels, but only rows of panels or contiguous partial areas of rows of panels can be replaced here. Also, a clearance free fastening of the panels without hindering their longitudinal changes caused by thermal effects is practically impossible in this manner.
In U.S. Pat. No. 5,575,439 an arrangement for removably securing a TPS panel to a space plane is seen. This arrangement employs three fastening points for the underside of each TPS panel with the aerostructure of a space plane. One point, which is provided at two opposing corners of a TPS panel, employs a fixed finger secured to the panel which slips under a fixed finger secured to the aerostructure. The second point, which is provided at one corner of each panel, employs a slotted member secured to the panel that slidingly engages an assembling bolt. The third or fixed point, that is provided on the corner of a panel opposed to the second point, includes a spring member that is secured to the aerostructure and which may be deformed by insertion of a tool from the exterior of the panel to deform the spring member sufficiently to permit it to capture a bolt carried by the panel upon release from engagement with the tool. No provision is made to isolate the TPS panels from substructure strains or to isolate the substructure from direct thermal conductance through the panel.
A removable thermal insulation blanket for mounting on the exterior of a reusable launch vehicle is disclosed in U.S. Pat. No. 5,928,752. This patent shows a blanket having first and second low density ceramic batting layers with a metal screen disposed between such layers. A plurality of fasteners are secured to the screen and are then further secured to mating fasteners positioned on the exterior surface of the launch vehicle by a means of a tool inserted directly through the outer batting layer. There is no provision to accommodate thermal expansion differences that will occur between a TPS panel and the aerosubstructure during such operations as fueling, ascent and reentry of the space vehicle.
In U.S. Pat. No. 5,489,074 an arrangement is disclosed for removably attaching TPS panels to the body of a space vehicle. This arrangement teaches a complex mechanism comprising a number of springs, latches and machined housings to isolate the movement of each panel. A wire that is permanently installed in the gap between two fairing elements of adjacent thermal insulation modules and which may be activated from outside the TPS panels to release a ball-locking bolt which secures the panel to the substructure and permit an insulative panel to be separated from the underlying insulative layer positioned directly upon the substructure of the space vehicle. While one TPS panel could apparently be removed by pulling the release wires on the four sides of the illustrated panel, it is not apparent how the panel could subsequently be replaced since the locking mechanism is buried beneath each side of each panel and is secured to the substructure of the space vehicle.
U.S. Pat. Nos. 5,675,950; 5,713,168; and 5,862,643 assigned to a common assignee all relate to a stand-off pedestal system for constructing a raised floor above an existing floor of a room for the purpose of providing a computer room or data processing center with sufficient space to accommodate cables, pipes, hoses, conduits and other routings for computer interconnections. The intent is to provide a constant standardized distance from the existing floor to the same size floor panels for easily affording access to the free space there below. This construction is intended to be a static design in an environmentally controlled computer room and is not intended to relieve thermal and structural deflections between thermal protection panels for a space vehicle.
As illustrated by the prior patents noted above, efforts have been made to provide an attachment means for securing thermal protection system panels to the aerostructure of a reusable launch vehicle or other space vehicle but such efforts have provided attachments that are heavy and complex and very time consuming in the installation of such panels. In addition, because of the interconnection of such panels and connecting arrangements it is very difficult or impossible to remove and replace one or a few separated TPS panels that may have been damaged in flight.
It is a general object of this invention to provide a strain compatible arrangement for providing support for a thermal protection panel system for a space vehicle.
It is another object of the present invention to provide a stand-off arrangement for a thermal protection panel system that can accommodate thermal expansion differences that occur during fueling of the space vehicle and during ascent and re-entry into the earth's atmosphere.
It is yet another object of the invention to support thermal protection panels at the outside mold line surface of a space vehicle.
It is a further object of the invention to provide a strain compatible stand-off arrangement for a thermal protection panel system that will permit ease of installation and ready removal and replacement of individual panels.
The foregoing has outlined some of the more pertinent objects of the invention. These objects should be construed to be merely illustrative of some of the more prominent features and applications of the intended invention. Many other beneficial results can be attained by applying the disclosed invention in a different manner or by modifying the invention withing the scope of the disclosure. Accordingly, other objects and a fuller understanding of the invention may be had by referring to the summary of the invention and the detailed description of the preferred embodiment in addition to the scope of the invention defined by the claims taken in conjunction with the accompanying drawings.